Frame connection between fan case and core housing in a gas turbine engine

ABSTRACT

A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A fan case surrounds the fan rotor. A core engine has a compressor section and includes a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of aft connecting members rigidly connected to the fan case, and to the core engine. A plurality of fan exit guide vanes are rigidly connected to the fan case, with the fan exit guide vanes including structural fan exit guide vanes which are rigidly connected to the core engine, and non-structural fan exit guide vanes, and the non-structural fan exit guide vanes being provided with an acoustic feature to reduce noise.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/407,275 filed Aug. 20, 2021.

BACKGROUND OF THE INVENTION

This application relates to incorporating frame connections and fan exitguide vanes connections between a fan case and a core housing in a gasturbine engine.

Gas turbine engines are known, and typically include a fan deliveringair into a bypass duct as propulsion air, and into a core enginehousing. The core engine housing houses a compressor section. The air iscompressed and delivered into a combustor where it is mixed with fueland ignited. Products of this combustion pass downstream over turbinerotors, driving them to rotate. The turbine rotors in turn rotate thefan and compressor rotors.

Historically the fan rotor was fixed to rotate at the same speed as afan drive turbine rotor, which may also drive a low pressure compressorrotor. More recently a gear reduction has been incorporated between thefan drive turbine and the fan rotor, allowing the fan rotor to rotate atslower speeds than the fan drive turbine.

In modern gas turbine engines with such a gear reduction the fan casehas been fixed to the core housing through a plurality of fan exit guidevanes, which provide the structural support between the fan case and theinner core housing.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine includes a fan rotordriven by a fan drive turbine about an axis through a gear reduction toreduce a speed of the fan rotor relative to a speed of the fan driveturbine. A fan case surrounds the fan rotor. A core engine has acompressor section and includes a low pressure compressor. The fan rotordelivers air into a bypass duct defined between the fan case and thecore engine. A rigid connection between the fan case and the core engineincludes a plurality of aft connecting members rigidly connected to thefan case, and to the core engine. A plurality of fan exit guide vanesare rigidly connected to the fan case, with the fan exit guide vanesincluding structural fan exit guide vanes which are rigidly connected tothe core engine, and non-structural fan exit guide vanes, and thenon-structural fan exit guide vanes being provided with an acousticfeature to reduce noise.

In another embodiment according to the previous embodiment, the acousticfeature includes the non-structural fan exit guide vanes are formed withchambers and a covering perforated face sheet.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes include 0-55% of a total number of fanexit guide vanes including the non-structural fan exit guide vane.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes include 15-30% of the total fan exitguide vanes.

In another embodiment according to any of the previous embodiments, thelow pressure compressor has four to six stages.

In another embodiment according to any of the previous embodiments, theaft connecting members are A-frames connecting the fan case to the coreengine.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes extend at a first angle from a radiallyinner end to a radially outer end. The first angle has a radiallyoutward component and an axially aft component. The A-frame legs extendfrom a radially inner connection to the core engine to a connectionpoint with the fan case at a second angle having a component in aradially outward direction, and in an axially forward direction.

In another embodiment according to any of the previous embodiments, thecore engine includes a fan intermediate case forward of the low pressurecompressor, and the structural fan guide vanes are rigidly connected tothe fan intermediate case.

In another embodiment according to any of the previous embodiments, thecore engine includes a compressor intermediate case intermediate the lowpressure compressor and a high pressure compressor, and the aftconnecting members rigidly secured to the compressor intermediate case.

In another embodiment according to any of the previous embodiments, thenon-structural fan exit guide vanes also are connected to the fanintermediate case, but being allowed to float in a radial direction.

In another featured embodiment, a gas turbine engine includes a fanrotor driven by a fan drive turbine about an axis through a gearreduction to reduce a speed of the fan rotor relative to a speed of thefan drive turbine. A fan case surrounds the fan rotor. A core engine hasa compressor section, including a low pressure compressor. The fan rotordelivers air into a bypass duct defined between the fan case and thecore engine. A rigid connection between the fan case and core engineincludes a plurality of A-frames each including a pair of legs rigidlyconnected at a connection point to the fan case. Each leg in the pairextend away from the connection point in opposed circumferentialdirections to be connected to the core engine to form an A-shape. Aplurality of fan exit guide vanes are rigidly connected to the fan case,with the fan exit guide vanes including structural fan exit guide vaneswhich are rigidly connected to the core engine, and non-structural fanexit guide vanes, and the non-structural fan exit guide vanes beingprovided with an acoustic feature to reduce noise.

In another embodiment according to any of the previous embodiments, theacoustic feature includes the non-structural fan exit guide vanes areformed with chambers and a covering perforated face sheet.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes include 0-55% of a total number of fanexit guide vanes including the non-structural fan exit guide vane.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes include 15-55% of the total fan exitguide vanes.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes include 15-30% of the total fan exitguide vanes.

In another embodiment according to any of the previous embodiments, thelow pressure compressor has four to six stages.

In another embodiment according to any of the previous embodiments,there are four of the A-frames connecting the fan case to the coreengine.

In another embodiment according to any of the previous embodiments, thestructural fan exit guide vanes extend at a first angle from a radiallyinner end to a radially outer end. The first angle has a radiallyoutward component and an axially aft component. The A-frame legs extendfrom a radially inner connection to the core engine to the connectionpoint with the fan case at a second angle having a component in aradially outward direction, and in an axially forward direction.

In another embodiment according to any of the previous embodiments, thecore engine includes a fan intermediate case forward of the low pressurecompressor, and the structural fan guide vanes are rigidly connected tothe fan intermediate case.

In another embodiment according to any of the previous embodiments, thecore engine includes a compressor intermediate case intermediate the lowpressure compressor and a high pressure compressor. The A-frames arerigidly secured to the compressor intermediate housing.

The present disclosure may include any one or more of the individualfeatures disclosed above and/or below alone or in any combinationthereof.

These and other features of the present invention can be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2A schematically shows details of the connection between a fan caseand an inner core housing.

FIG. 2B shows another connection detail.

FIG. 3A shows a simplified rear view of FIG. 2 .

FIG. 3B shows a detail of a plurality of fan exit guide vanes.

FIG. 3C shows another embodiment.

FIG. 3D shows another embodiment.

FIG. 4 shows details of one fan exit guide vane.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 mayinclude a single-stage fan 42 having a plurality of fan blades 43. Thefan blades 43 may have a fixed stagger angle or may have a variablepitch to direct incoming airflow from an engine inlet. The fan 42 drivesair along a bypass flow path B in a bypass duct 13 defined within ahousing 15 such as a fan case or nacelle, and also drives air along acore flow path C for compression and communication into the combustorsection 26 then expansion through the turbine section 28. A splitter 29aft of the fan 42 divides the air between the bypass flow path B and thecore flow path C. The housing 15 may surround the fan 42 to establish anouter diameter of the bypass duct 13. The splitter 29 may establish aninner diameter of the bypass duct 13. Although depicted as a two-spoolturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with two-spool turbofans as the teachings may be applied to othertypes of turbine engines including three-spool architectures. The engine20 may incorporate a variable area nozzle for varying an exit area ofthe bypass flow path B and/or a thrust reverser for generating reversethrust.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in the exemplary gas turbineengine 20 is illustrated as a geared architecture 48 to drive the fan 42at a lower speed than the low speed spool 30. The inner shaft 40 mayinterconnect the low pressure compressor 44 and low pressure turbine 46such that the low pressure compressor 44 and low pressure turbine 46 arerotatable at a common speed and in a common direction. In otherembodiments, the low pressure turbine 46 drives both the fan 42 and lowpressure compressor 44 through the geared architecture 48 such that thefan 42 and low pressure compressor 44 are rotatable at a common speed.Although this application discloses geared architecture 48, its teachingmay benefit direct drive engines having no geared architecture. The highspeed spool 32 includes an outer shaft 50 that interconnects a second(or high) pressure compressor 52 and a second (or high) pressure turbine54. A combustor 56 is arranged in the exemplary gas turbine 20 betweenthe high pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 57 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

Airflow in the core flow path C is compressed by the low pressurecompressor 44 then the high pressure compressor 52, mixed and burnedwith fuel in the combustor 56, then expanded through the high pressureturbine 54 and low pressure turbine 46. The mid-turbine frame 57includes airfoils 59 which are in the core flow path C. The turbines 46,54 rotationally drive the respective low speed spool 30 and high speedspool 32 in response to the expansion. It will be appreciated that eachof the positions of the fan section 22, compressor section 24, combustorsection 26, turbine section 28, and fan drive gear system 48 may bevaried. For example, gear system 48 may be located aft of the lowpressure compressor, or aft of the combustor section 26 or even aft ofturbine section 28, and fan 42 may be positioned forward or aft of thelocation of gear system 48.

The fan 42 may have at least 10 fan blades 43 but no more than 20 or 24fan blades 43. In examples, the fan 42 may have between 12 and 18 fanblades 43, such as 14 fan blades 43. An exemplary fan size measurementis a maximum radius between the tips of the fan blades 43 and the enginecentral longitudinal axis A. The maximum radius of the fan blades 43 canbe at least 38 inches, or more narrowly no more than 75 inches. Forexample, the maximum radius of the fan blades 43 can be between 45inches and 60 inches, such as between 50 inches and 55 inches. Anotherexemplary fan size measurement is a hub radius, which is defined asdistance between a hub of the fan 42 at a location of the leading edgesof the fan blades 43 and the engine central longitudinal axis A. The fanblades 43 may establish a fan hub-to-tip ratio, which is defined as aratio of the hub radius divided by the maximum radius of the fan 42. Thefan hub-to-tip ratio can be less than or equal to 0.35, or more narrowlygreater than or equal to 0.20, such as between 0.25 and 0.30. Thecombination of fan blade counts and fan hub-to-tip ratios disclosedherein can provide the engine 20 with a relatively compact fanarrangement.

The low pressure compressor 44, high pressure compressor 52, highpressure turbine 54 and low pressure turbine 46 each include one or morestages having a row of rotatable airfoils. Each stage may include a rowof vanes adjacent the rotatable airfoils. The rotatable airfoils areschematically indicated at 47, and the vanes are schematically indicatedat 49.

The low pressure compressor 44 and low pressure turbine 46 can includean equal number of stages. For example, the engine 20 can include athree-stage low pressure compressor 44, an eight-stage high pressurecompressor 52, a two-stage high pressure turbine 54, and a three-stagelow pressure turbine 46 to provide a total of sixteen stages. In otherexamples, the low pressure compressor 44 includes a different (e.g.,greater) number of stages than the low pressure turbine 46. For example,the engine 20 can include a five-stage low pressure compressor 44, anine-stage high pressure compressor 52, a two-stage high pressureturbine 54, and a four-stage low pressure turbine 46 to provide a totalof twenty stages. In other embodiments, the engine 20 includes afour-stage low pressure compressor 44, a nine-stage high pressurecompressor 52, a two-stage high pressure turbine 54, and a three-stagelow pressure turbine 46 to provide a total of eighteen stages. It shouldbe understood that the engine 20 can incorporate other compressor andturbine stage counts, including any combination of stages disclosedherein.

The engine 20 may be a high-bypass geared aircraft engine. The bypassratio can be greater than or equal to 10.0 and less than or equal toabout 18.0, or more narrowly can be less than or equal to 16.0. Thegeared architecture 48 may be an epicyclic gear train, such as aplanetary gear system or a star gear system. The epicyclic gear trainmay include a sun gear, a ring gear, a plurality of intermediate gearsmeshing with the sun gear and ring gear, and a carrier that supports theintermediate gears. The sun gear may provide an input to the gear train.The ring gear (e.g., star gear system) or carrier (e.g., planetary gearsystem) may provide an output of the gear train to drive the fan 42. Agear reduction ratio may be greater than or equal to 2.3, or morenarrowly greater than or equal to 3.0, and in some embodiments the gearreduction ratio is greater than or equal to 3.4. The gear reductionratio may be less than or equal to 4.0. The fan diameter issignificantly larger than that of the low pressure compressor 44. Thelow pressure turbine 46 can have a pressure ratio that is greater thanor equal to 8.0 and in some embodiments is greater than or equal to10.0. The low pressure turbine pressure ratio can be less than or equalto 13.0, or more narrowly less than or equal to 12.0. Low pressureturbine 46 pressure ratio is pressure measured prior to an inlet of lowpressure turbine 46 as related to the pressure at the outlet of the lowpressure turbine 46 prior to an exhaust nozzle. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans. All of these parameters are measured at the cruise conditiondescribed below.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. The engine parameters described above, and those in thenext paragraph are measured at this condition unless otherwisespecified.

“Fan pressure ratio” is the pressure ratio across the fan blade 43alone, without a Fan Exit Guide Vane (“FEGV”) system. A distance isestablished in a radial direction between the inner and outer diametersof the bypass duct 13 at an axial position corresponding to a leadingedge of the splitter 29 relative to the engine central longitudinal axisA. The fan pressure ratio is a spanwise average of the pressure ratiosmeasured across the fan blade 43 alone over radial positionscorresponding to the distance. The fan pressure ratio can be less thanor equal to 1.45, or more narrowly greater than or equal to 1.25, suchas between 1.30 and 1.40. “Corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram ° R)/(518.7° R)]^(0.5). The corrected fan tip speedcan be less than or equal to 1150.0 ft/second (350.5 meters/second), andcan be greater than or equal to 1000.0 ft/second (304.8 meters/second).

The fan 42, low pressure compressor 44 and high pressure compressor 52can provide different amounts of compression of the incoming airflowthat is delivered downstream to the turbine section 28 and cooperate toestablish an overall pressure ratio (OPR). The OPR is a product of thefan pressure ratio across a root (i.e., 0% span) of the fan blade 43alone, a pressure ratio across the low pressure compressor 44 and apressure ratio across the high pressure compressor 52. The pressureratio of the low pressure compressor 44 is measured as the pressure atthe exit of the low pressure compressor 44 divided by the pressure atthe inlet of the low pressure compressor 44. In examples, a product ofthe pressure ratio of the low pressure compressor 44 and the fanpressure ratio is between 3.0 and 6.0, or more narrowly is between 4.0and 5.5. The pressure ratio of the high pressure compressor ratio 52 ismeasured as the pressure at the exit of the high pressure compressor 52divided by the pressure at the inlet of the high pressure compressor 52.In examples, the pressure ratio of the high pressure compressor 52 isbetween 7.0 and 12.0, or more narrowly is between 10.0 and 11.5. The OPRcan be equal to or greater than 44.0, and can be less than or equal to70.0, such as between 50.0 and 60.0. The overall and compressor pressureratios disclosed herein are measured at the cruise condition describedabove, and can be utilized in two-spool architectures such as the engine20 as well as three-spool engine architectures.

The engine 20 establishes a turbine entry temperature (TET). The TET isdefined as a maximum temperature of combustion products communicated toan inlet of the turbine section 28 at a maximum takeoff (MTO) condition.The inlet is established at the leading edges of the axially forwardmostrow of airfoils of the turbine section 28, and MTO is measured atmaximum thrust of the engine 20 at static sea-level and 86 degreesFahrenheit (° F.). The TET may be greater than or equal to 2700.0° F.,or more narrowly less than or equal to 3500.0° F., such as between2750.0° F. and 3350.0° F. The relatively high TET can be utilized incombination with the other techniques disclosed herein to provide acompact turbine arrangement.

The engine 20 establishes an exhaust gas temperature (EGT). The EGT isdefined as a maximum temperature of combustion products in the core flowpath C communicated to at the trailing edges of the axially aftmost rowof airfoils of the turbine section 28 at the MTO condition. The EGT maybe less than or equal to 1000.0° F., or more narrowly greater than orequal to 800.0° F., such as between 900.0° F. and 975.0° F. Therelatively low EGT can be utilized in combination with the othertechniques disclosed herein to reduce fuel consumption.

FIG. 2A shows an engine 100, which may be similar to the engine 20 ofFIG. 1 . A shaft 102 is driven by a fan drive turbine to drive a fanrotor 106 through a gear reduction 104. The drive connection here may begenerally as described above with regard to FIG. 1 . A fan case 108surrounds the fan rotor 106, and an core engine 400 may include asplitter wall 110 that surrounds compressor housing wall 111 whichhouses a low pressure compressor 112, and a high pressure compressor230, and combustor and turbine sections (not shown in this Figure). Thecore engine 400 must be rigidly connected to the fan case 108, toaddress torque and other loads.

Applicant has previously developed a geared gas turbine engine. In thisfirst generation engine the fan case was connected to the core enginethrough a plurality of fan exit guide vanes. Each of these fan exitguide vanes were structural elements that provided a load path betweenthe fan case and the core engine.

In engine 100, as will be described below, there are fewer structuralfan exit guide vanes 114. A-frames 116 have been added to provideadditional rigidity.

As shown, the low pressure compressor 112 has five rotating stages. Inembodiments the low pressure compressor may have four to six stages,which is longer than the first generation gas turbine enginemanufactured by Applicant mentioned above. With such a long low pressurecompressor 112, mounting challenges are raised.

As can been seen from FIGS. 2A and 2B, the fan exit guide vanes 114/129extend from an inner point 200 attached at an angle to a radially outerpoint 201, with the angle having a component in a radially outerdirection, and another component in an axially aft direction.

Conversely, the A-frames 116 extend from a radially inner point 202attached to the core engine 400, radially outwardly at an angle to anouter point 203 connected to the fan case 108. The angle of the A-frame116 has a component in a radially outer direction and another componentin an axially forward direction.

As shown in FIG. 2A, the core engine 400 includes a fan intermediatecase 210 including a plurality of struts, and having a mount bracket220. A non-structural guide vane 129 is attached to the bracket 220through pins 222. In this manner, the non-structural guide vanes 129 can“float” or adjust radially relative to the bracket 220, but areprevented from moving circumferentially. Alternatively, thenon-structural guide vanes may be fixed to bracket 220.

FIG. 2B shows a detail of the mount of a structural guide vane 114 tothe fan intermediate case 210 and to the bracket 220 through a first pin300 preventing circumferential movement, and a second pin 302 preventingradial movement.

Returning to FIG. 2A, it can be seen that the A-frames 116 are attachedat inner ends 202 at a bracket 234 which is fixed with a compressorintermediate case 232 having a plurality of struts. This viewillustrates one strut. The compressor intermediate cases 232 isintermediate the low pressure compressor 112 and a high pressurecompressor 230.

Although specific mount locations are shown, other connections betweenthe fan exit guide vanes 114/129 and A-frames 116 to the core engine 400may be utilized. For purposes of this application, the core engine isdefined to include at least the compressor housing wall 111, the fanintermediate case 210, the compressor intermediate case 232, the lowpressure compressor 112, the high pressure compressor 230 and acombustor end turbine section, not shown, but which may be as disclosedwith regard to FIG. 1 .

As shown in FIG. 3A, the A-frames 116 comprise two rigid members 118 and120 which extend from a connection point 122 at the fan case 108inwardly to connection points 124 with the compressor housing 111. Ascan be seen, the members 118 extend away from each other moving awayfrom the connection point 122, and in opposed directions/angles relativeto a plane drawn through connection point 122, parallel to a center axisX. In FIG. 3A there are sixteen structural guide vanes 114 illustrated.Other numbers can be used. They are rigidly connected at 126 to the fancase 108, and at 128. It should be understood that FIG. 3A is an oversimplification. In fact, as mentioned above, the fan exit guide vanes114 and 129 and A-frames 116 may be attached as shown in FIGS. 2A and2B. The view of FIG. 3A is shown simply to illustrate some generalrelationships. In embodiments, the engine 100 can have more or less thaneight of the A-frame legs 118 and 120.

As shown in FIG. 3B, there are forty-eight fan exit guide vanes total,with sixteen of the fan exit guide vanes being structural vanes 114.Intermediate each structural fan exit guide vane 114 are twonon-structural guide vanes 129. The non-structural guide vanes 129 arerigidly connected to the fan case 108 at 130, but as mentioned may floatrelative to the wall 110 at a radially inner point 132. Alternatively,inner point 132 may be fixed.

FIG. 3C schematically shows an embodiment 150 having three A-frames 152connecting the fan case to the core engine.

FIG. 3D shows an embodiment 160 having five A-frames 162 connecting thefan case to the core engine. As can appreciated, other numbers ofA-frames may be utilized.

In fact, rigid connections other than A-frames may be utilized at thisaft location. For purposes of interpreting this application, theA-frames along with the other types of rigid connections may bedescribed generically as aft connecting numbers.

Since the guide vanes 129 are not structural, they can provide anacoustic function. As shown in FIG. 4 , one of the non-structural fanexit guide vanes 129 is illustrated. The guide vane has a pressure wall134 and a suction wall 136. An outer skin 137 is perforated at 138 andprovided over a plurality of chambers 140. The chambers 140 can have anynumber of shapes including honeycomb, or other cross-sections. Theperforations 138 could be circular, but can be other shapes, includingelongated slots.

In embodiments the structural guide vanes 114 may include 0% to 55% ofthe total fan exit guide vanes. In other embodiments the structural fanexit guide vanes 114 may be 15 to 55% of the total fan exit guide vanes.In other embodiments the structural fan exit guide vanes 114 may include15 to 50% of the total fan exit guide vanes. In yet another embodiment,the structural fan exit guide vanes provide 15 to 30% of the total fanexit guide vanes.

A gas turbine engine under this disclosure could be said to include afan rotor driven by a fan drive turbine about an axis through a gearreduction to reduce a speed of the fan rotor relative to a speed of thefan drive turbine. A fan case surrounds the fan rotor, and a core enginehas a compressor section, including a low pressure compressor. The fanrotor delivers air into a bypass duct defined between the fan case andthe core engine. A rigid connection between the fan case and the coreengine includes a plurality of aft connecting members rigidly connectedto the fan case, and to the core engine. A plurality of fan exit guidevanes are rigidly connected to the fan case, with the fan exit guidevanes including structural fan exit guide vanes which are rigidlyconnected to the core engine, and non-structural fan exit guide vanes,and the non-structural fan exit guide vanes are provided with anacoustic feature to reduce noise.

A gas turbine engine under this disclosure could also be said to includea fan rotor driven by a fan drive turbine about an axis through a gearreduction to reduce a speed of the fan rotor relative to a speed of thefan drive turbine. A fan case surrounds the fan rotor. A core engine hasa compressor section includes a low pressure compressor. The fan rotordelivers air into a bypass duct defined between the fan case and thecore engine. A rigid connection between the fan case and inner coreengine includes a plurality of A-frames each including a pair of legsrigidly connected at a connection point to the fan case. Each leg in thepair extending away from the connection point in opposed circumferentialdirections to be connected to the core engine to form an A-shape. Aplurality of fan exit guide vanes are rigidly connected to the fan case,with the fan exit guide vanes including structural fan exit guide vaneswhich are rigidly connected to the core engine, and non-structural fanexit guide vanes, and the non-structural fan exit guide vanes areprovided with an acoustic feature to reduce noise.

Although embodiments of this disclosure have been shown, a worker ofordinary skill in this art would recognize that modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the true scope and content of thisdisclosure.

What is claimed is:
 1. A gas turbine engine comprising: a fan rotordriven by a fan drive turbine about an axis through a gear reduction toreduce a speed of said fan rotor relative to a speed of said fan driveturbine; a fan case surrounding said fan rotor, and a core engine with acompressor section, including a low pressure compressor, said lowpressure compressor having four to six stages; said fan rotor deliveringair into a bypass duct defined between said fan case and said coreengine, and a rigid connection between said fan case and said coreengine including a plurality of aft connecting members rigidly connectedto said fan case, and to said core engine at a location downstream ofsaid low pressure compressor; a plurality of fan exit guide vanesrigidly connected to said fan case, with said plurality of fan exitguide vanes including structural fan exit guide vanes which are rigidlyconnected to said core engine and non-structural fan exit guide vanes;and said structural fan exit guide vanes include more than 0% and lessthan or equal to 55% of the plurality of fan exit guide vanes includingthe non-structural fan exit guide vane.
 2. The gas turbine engine as setforth in claim 1, wherein said non-structural fan exit guide vanesinclude an acoustic feature.
 3. The gas turbine engine as set forth inclaim 1, wherein at least one non-structural fan exit guide vane of thenon-structural fan exit guide vanes is formed with chambers configuredto reduce noise associated with the gas turbine engine.
 4. The gasturbine engine as set forth in claim 3, wherein at least onenon-structural fan exit guide vane of the non-structural fan exit guidevanes includes a covering comprising a perforated face sheet.
 5. The gasturbine engine as set forth in claim 1, wherein at least onenon-structural fan exit guide vane of the non-structural fan exit guidevanes includes a covering comprising a perforated face sheet.
 6. The gasturbine engine as set forth in claim 1, wherein said structural fan exitguide vanes include 15-30% of the plurality of fan exit guide vanes. 7.The gas turbine engine as set forth in claim 1, wherein said aftconnecting members are A-frames connecting said fan case to said coreengine.
 8. The gas turbine engine as set forth in claim 7, wherein saidstructural fan exit guide vanes extend at a first angle from a radiallyinner end to a radially outer end, and said first angle having aradially outward component and an axially aft component, and saidA-frame legs extending from a radially inner connection to said coreengine to a connection point with said fan case at a second angle havinga component in a radially outward direction, and in an axially forwarddirection.
 9. The gas turbine engine as set forth in claim 1, whereinsaid core engine includes a fan intermediate case forward of said lowpressure compressor, and said structural fan guide vanes are rigidlyconnected to said fan intermediate case.
 10. The gas turbine engine asset forth in claim 1, wherein said core engine including a compressorintermediate case intermediate said low pressure compressor and a highpressure compressor, and said aft connecting members rigidly secured tosaid compressor intermediate case.
 11. The gas turbine engine as setforth in claim 1, wherein a bypass ratio defined as the volume of airdelivered by the fan rotor into the bypass duct divided by the volume ofair delivered by the fan rotor into the core engine, and said bypassratio being greater than or equal to 10.0 and less than or equal to18.0.
 12. The gas turbine engine as set forth in claim 1, wherein thefan rotor having a fan hub with fan blades extending radially outwardlyof the fan hub, and there being a hub radius defined as a distancebetween the fan hub at a location of a leading edge of the fan blade andan engine rotational axis, and there being a fan hub to fan blade tipratio defined as the ratio of the hub radius divided by a maximum radiusof the fan, and the fan hub to fan blade tip ratio being greater than orequal 0.20 and less than or equal to 0.35.
 13. A gas turbine enginecomprising: a fan rotor driven by a fan drive turbine about an axisthrough a gear reduction to reduce a speed of said fan rotor relative toa speed of said fan drive turbine; a fan case surrounding said fanrotor, and a core engine with a compressor section, including a lowpressure compressor; said fan rotor delivering air into a bypass ductdefined between said fan case and said core engine, and a rigidconnection between said fan case and core engine including a pluralityof A-frames each including a pair of legs rigidly connected at aconnection point to said fan case, and each leg in said pair extendingaway from said connection point in opposed circumferential directions tobe connected to said core engine to form an A-shape; and a plurality offan exit guide vanes rigidly connected to said fan case, with saidplurality of fan exit guide vanes including structural fan exit guidevanes which are rigidly connected to said core engine, andnon-structural fan exit guide vanes, and said non-structural fan exitguide vanes being provided with an acoustic feature to reduce noise,said acoustic feature including said non-structural fan exit guide vanesbeing formed with chambers and a covering perforated face sheet.
 14. Thegas turbine engine as set forth in claim 13, wherein the plurality ofA-frames includes between three and five A-frames connecting said fancase to said core engine.
 15. The gas turbine engine as set forth inclaim 13, wherein said structural fan exit guide vanes extend at a firstangle from a radially inner end to a radially outer end, and said firstangle having a radially outward component and an axially aft component,and said A-frame legs extending from a radially inner connection to saidcore engine to said connection point with said fan case at a secondangle having a component in a radially outward direction, and in anaxially forward direction.
 16. The gas turbine engine as set forth inclaim 13, wherein said core engine includes a fan intermediate caseforward of said low pressure compressor, and at least one structural fanguide vane of said structural fan guide vanes is rigidly connected tosaid fan intermediate case.
 17. The gas turbine engine as set forth inclaim 13, wherein said core engine including a compressor intermediatecase intermediate said low pressure compressor and a high pressurecompressor, and at least one A-frame of said plurality of A-frames beingrigidly secured to said compressor intermediate housing.
 18. The gasturbine engine as set forth in claim 13, wherein a bypass ratio definedas the volume of air delivered by the fan rotor into the bypass ductdivided by the volume of air delivered by the fan rotor into the coreengine, and said bypass ratio being greater than or equal to 10.0 andless than or equal to 18.0.
 19. The gas turbine engine as set forth inclaim 13, wherein the fan rotor having a fan hub with fan bladesextending radially outwardly of the fan hub, and there being a hubradius defined as a distance between the fan hub at a location of aleading edge of the fan blade and an engine rotational axis, and therebeing a fan hub to fan blade tip ratio defined as the ratio of the hubradius divided by a maximum radius of the fan, and the fan hub to fanblade tip ratio being greater than or equal 0.20 and less than or equalto 0.35.
 20. The gas turbine engine as set forth in claim 13, whereinsaid low pressure compressor has four to six stages.